1. Field of the Invention
The present invention relates to a gas turbine stationary blade, and more specifically to a gas turbine stationary blade having a cooling structure for applying air cooling to a second stage stationary blade with a high cooling efficiency.
2. Description of the Prior Art
In FIG. 1, a cross sectional view of a typical structure of gas turbine is shown and an outline thereof will be described first. In FIG. 1, numeral 1 designates a compressor portion, numeral 2 designates a combustor portion and numeral 3 designates a turbine portion. Numeral 4 designates a rotor, which extends in a turbine axial direction from the compressor portion 1 to the turbine portion 3 . Numeral 6 designates an inner housing and numerals 7, 8 designate cylinders of compressor portion 1, which surround an outer circumference of a compressor. Numeral 9 designates a cylindrical shell forming a chamber, numeral 10 designates an outer shell of the turbine portion 3, numeral 11 designates an inner shell of the turbine portion 3 numeral 12 designates a stationary blade of the compressor, and a plurality of the stationary blades being disposed along a compressor circumferential direction with equal spacing between each of the blades and in multi-stages along a compressor axial direction, and numeral 13 designates a moving blade of the compressor, a plurality of the moving blades being fixed around the rotor 4 and disposed alternately with the stationary blades 12 along the compressor axial direction.
Numeral 14 designates a chamber surrounded by the cylindrical shell 9 and numeral 15 designates a combustor, disposed in the chamber 14, into which fuel 35 is injected from a fuel nozzle 34 for combustion. Numeral 16 designates a duct for leading a high temperature combustion gas 30 generated in the combustor 15 into the turbine portion 3. Numeral 17 designates a second stage stationary blade of the gas turbine, which is the object of the present invention. In the case shown in FIG. 1, the gas turbine is constructed of four stage stationary blades and four stage moving blades disposed alternately therewith, and the high temperature combustion gas 30 passes through the blades and is discharged as an expanded gas 3. Numeral 21 designates a manifold of the compressor portion 1 and numeral 22 designates a manifold of the turbine portion 3. Cooling air is supplied from the manifold 21 of the compressor portion 1 to the manifold 22 of the turbine portion via a pipe 32 and an air piping 19.
In the gas turbine constructed as mentioned above, the fuel 35 is injected into the combustor 15 from the fuel nozzle 34 to be burnt to generate the high temperature combustion gas 30 and then flows into the turbine portion 3 to pass through a passage where the stationary blades and the moving blades are disposed alternately and to expand to rotate the moving blades and the rotor 4 and is discharged as the expanded gas 31.
On the other hand, while a portion of the cooling air is supplied from the compressor portion into the moving blades of the gas turbine for cooling thereof via rotor discs, a portion of the cooling air is also supplied from the manifold 21 of the compressor portion 1 into the manifold 22 of the turbine portion 3 for cooling of the second stage stationary blade 17 as well as to be used as seal air via pipe 32 and the air piping 19.
Next, the second stage stationary blade 17 will be described in detail. FIG. 6 is a cross sectional view of the second stage stationary blade 17 of the prior art gas turbine, the stationary blade being cut along a turbine axial direction at approximately a central portion of its inner shroud and seen from an inner side thereof, that is, on a rotor 4 side. FIG. 7 is a cross sectional view taken on line D--D of FIG. 6, FIG. 8 is a cross sectional view taken on line E--E of FIG. 6, FIG. 9 is a cross sectional view taken on line F--F of FIG. 6, FIG. 10 is a cross sectional view taken on line G--G of FIG. 6, FIG. 11 is a cross sectional view taken on line H--H of FIG. 6 and FIG. 12 is a cross sectional view taken on line J--J of FIG. 6.
In FIG. 6, numeral 26 designates an inner shroud and provided therein are a rib 40, a leading edge passage 42 and a trailing edge passage 44 mutually separated by the rib 40, and a projection portion 95 provided therearound. Numerals 96, 97 designate rails of both side edge portions of the inner shroud 26 and numerals 93, 94 designate passages of cooling air provided in the rails 96, 97, respectively. A passage 88 is provided in a leading edge portion 41 of the inner shroud 26 and a multiplicity of passages 92 are provided in a trailing edge portion 43 of the inner shroud 26. There are provided a multiplicity of needle-like fins in the passage 88, so that convection is accelerated and heat transfer efficiency is enhanced. Numeral 100 designates a recess portion formed by the projection portion 95 and numerals 83, 84 designate impingement plates, each having a multiplicity of small holes 101 provided therein as passages of air. Numerals 81, 82 designate a front flange and a rear flange, respectively, and there are provided passages 90, 91 in the front flange 81. Cooling air 57 which has entered the recess portion 100 passes through the passage 90 in the front flange 81 and the passage 88 in the leading edge portion 41 and then through the passage 91 in the front flange 81 and enters a chamber formed by the impingement plate 83. Also, a portion 58 of the cooling air which has entered the passage 88 passes through the passages 93, 94 in the rails 96, 97 of the side edge portions for cooling therearound and is discharged outside as a cooling air 61. The cooling air which has flowed through the small holes 101 of the impingement plates 83, 84 and the cooling air which has flown through the passage 91 gather together in the chamber to further flow through the multiplicity of passages 92 of the trailing edge portion 43 and to be discharged outside as a cooling air 60.
In FIG. 7, being a cross sectional view taken on line D--D of FIG. 6, the passage 88 is formed in the leading edge portion 41 of the inner shroud 26 and the multiplicity of needle-like fins 89 are provided therein. In a space between the front flange 81 and the rear flange 82 are provided the recess portion 100 in front of the projection portion 95 and a recess portion 99 to the rear of the projection portion 95. The impingement plate 84 is provided so as to form chamber 78 on an outer side of the impingement plate 84. In the front flange 81, there is provided on the passage 90 which connects to the passage 88. The portion 57 of the cooling air, flowing through the passages 90 and 88, and another portion 59 of the cooling air, passing through the small holes 101 of the impingement plate 84, gather together in the chamber 78 to further flow through the multiplicity of passages 92 of the trailing edge portion 43 and is then discharged as the cooling air 60.
In FIG. 8, being a cross sectional view taken on line E--E of FIG. 6, the second stage stationary blade 17 has the inner shroud 26 and the outer shroud 27 and a blade portion 25 is formed therebetween. The leading edge passage 42 in front of the rib 40 and the trailing edge passage 44 in the rear are formed between a leading edge portion 28 and a trailing edge portion 29 of the blade portion 25, and cylindrical members 46, 47 are inserted into these passages 42,44, respectively. There are provided a multiplicity of cooling air holes 70, 71 in side walls of the cylindrical members 46, 47, respectively, and also cooling air holes 72, 73 in bottom walls of the cylindrical members 46, 47, respectively. Further, there are provided a multiplicity of pins 62 in the trailing edge portion 29.
In the leading edge portion 41 of the inner shroud 26, the passage 88 and the needle-like fins 89 in the passage 88 are provided, and in the trailing edge portion 43 of the inner shroud 26, the passages 92 are provided so as to connect to a cavity 45 which is formed by the front and rear flanges 81, 82 and a seal support portion 66. A chamber 77 is formed by the impingement plate 84 in the cavity 45. On the inner side of the cavity 45, the seal support portion 66 supports a seal 33, by which a seal mechanism between the inner shroud 26 and rotor side arm portions 48 is constructed.
Cooling air 19' from the air piping 19 flows into the cylindrical members 46, 47 to be injected through the cooling air holes 70, 71 to impinge on walls of the leading edge passage 42 and the trailing edge passage 44 and to flow toward the inner side thereof as well as to be injected through the cooling air holes 72, 73 of the bottom walls of the cylindrical members 46, 47 to flow into opening portions 68, 69. Then the cooling air, as shown by numerals 75, 76 flows into the cavity 45. The cooling air then flows into a space between the inner shroud 26 and a front stage moving blade thereof and a space between the inner shroud 26 and a rear stage moving blade thereof via the seal 33 to thereby maintain the spaces in a higher pressure than in a passage of the high temperature combustion gas 30 to prevent the high temperature combustion gas 30 from coming into the spaces.
In FIG. 9, being a cross sectional view taken on line F--F of FIG. 6, a recess portion 98 and the chamber 77 are formed by the impingement plate 83 between the front flange 81 and the rear flange 82, and the passage 91 provided in the front flange 81 connects to the passage 88 and the passages 92 provided in the trailing edge portion 43 connect to the chamber 77. Cooling air 59 in the cavity 45 is injected into the chamber 77 through the small holes 101 of the impingement plate 83 for cooling therearound, as shown by arrows of the air 59. On the other hand, cooling air which has flowed through the passage 88 enters the passage 91 of the front flange 81 to join with the cooling air 59 in the chamber 77 so both are then discharged as the cooling air 60 through the passages 92 of the trailing edge portion 43.
In FIG. 10, being a cross sectional view taken on line G--G of FIG. 6, the recess portions 98, 99 are provided around the blade portion 25 and the passages 93, 94 are provided in the rails 96, 97, respectively. Also, the chambers 77, 78 are formed by the impingement plates 83, 84, respectively. Cooling air 75 flows into the cavity 45 from the leading edge passage 42 and flows therefrom into the chambers 77, 78 through the small holes 101 of the impingement plates 83, 84.
In FIG. 11, being a cross sectional view taken on line H--H of FIG. 6, the passages 90, 91 of the front flange 81 and the passages 93, 94 of the side edge portions are provided in both of the side edge portions of the inner shroud 26 and the passages 90, 91 connect to the passage 88 of the leading edge portion 41.
In FIG. 12, being a cross sectional view taken on line J--J of FIG. 6, the passage 94 of the rail 97 is provided extending through the trailing edge portion 43 so that the cooling air 61 is discharged therefrom and the impingement plate 83 is provided between the front flange 81 and the rear flange 82.
In the second stage stationary blade of a gas turbine described as above, the cooling air 57 from the recess portion 100 flows into the passage 88 of the leading edge portion 41 through the passage 90 of the front flange 81. There are provided the multiplicity of needle-like fins 89 in the passage 88, and thereby the cooling effect of the cooling air 57 is enhanced so that portions therearound are cooled efficiently. Then, the cooling air 57 bends approximately orthogonally at the passage 91 and flows into the chamber 77 formed by the impingement plate 83 to join with the cooling air flowing thereinto through the small holes 101 of the impingement plate 83 and flows together through the trailing edge portion 43 for cooling thereof and is discharged through the passages 92. Also, the cooling air which has been injected through the small holes 101 of the impingement plate 84 to enter the chamber 78 is likewise discharged through the passages 92.
Further, the portion 58 of the air which has entered the passage 88 passes through the passages 93, 94 in the rails 96, 97, respectively, of the side edge portions for cooling therearound and are discharged as the cooling air 61 from the trailing edge portion 43. Thus, the cooling air 75, 76 in the cavity 45 is portioned to be made effective use of, respectively flowing through the passage 88, in which heat transfer is enhanced by the needle-like fins 89, the passages 93, 94 in the rails 96, 97 and the multiplicity of passages 92 in the trailing edge portion 43, and thereby the entire cooling of the inner shroud 26 is aimed to be performed efficiently.
That is, according to the air cooled system of the second stage stationary blade of a gas turbine in the prior art as described above, in order to ensure the entire cooling effect of the inner shroud 26, the cooling air passes through the passage 88 and the needle-like fins 89 provided therein for enhancement of the cooling effect to further flow portionally into the chamber 77 formed by the impingement plate 83 through the passage 91 of the front flange 81, and also the cooling air is injected into the chambers 77, 78 through the small holes 101 of the impingement plates 83 , 84 for cooling of the central portion, and then both of the cooling air flows join together to flow through the multiplicity of passages 92 of the trailing edge portion 43 for cooling therearound. Further, the cooling air from the passage 88 of the leading edge portion 41 portionally flows through the passages 93, 94 of the rails 96, 97 of the side edge portions for cooling therearound.
According to the cooling structure mentioned above, however, while the entire inner shroud is cooled efficiently, the cooling air which has entered the passage 88 portionally flows out of the passage 91 for cooling of the central portion, hence in the leading edge portion 41 and the side edge portions which are especially exposed to the high temperature combustion gas, there occurs a shortage of the cooling air flowing in the passages 93, 94 of the side edge portions, resulting in insufficiency of cooling in the side edge portions.
Also, the cooling air entering the passage 88 of the leading edge portion 41 is a part of the cooling air entering the cavity 45 and comes from the recess portion 100 through the passage 90, and in order to further enhance the cooling effect of the leading edge portion 41, it is expected that the amount of the cooling air flowing therein and the flow velocity thereof are increased so as to enhance the cooling effect further.